1. Field of the Invention
The present invention relates to a turbojet engine capable of operating effectively at both supersonic and subsonic velocities, and to a method of operating such an engine.
2. Brief Description of the Prior Art
There has long been a requirement for an aircraft which is capable of operating effectively in the supersonic range, and yet have the capability of cruising subsonically with a relatively low specific fuel consumption. With regard to military aircraft, it is necessary that the aircraft be capable of developing very high thrusts for acceleration and operation in the high supersonic range. Yet, for many mission requirements, such as remaining aloft for long periods of time or traveling to and from the areas where the mission is to be carried out, it is desirable that the aircraft be capable of extended subsonic cruise with low specific fuel consumption. With regard to supersonic jet transport, the aircraft must of course be capable of efficient operation at supersonic cruise. However, for extended flight over populated land masses, where the aircraft is required to travel subsonically to avoid the effect of the supersonic boom over such populated areas, the aircraft should also have the capability of operating with low specific fuel consumption.
The problem arises in that when an engine has its operating components matched with one another to operate effectively at high power settings for acceleration and supersonic cruise, the same matching of engine components does not necessarily lend itself for efficient operation at subsonic cruise. At very high thrust settings, the engine necessarily burns fuel at a relatively high rate of consumption and developes combustion exit temperatures in the range of 2800.degree. F. which are forecast for commercial airplanes in the next decade. This necessarily means a relatively high corrected flow through the turbine, and the compressor and turbine must be matched so that the turbine can accept this high corrected flow. However, for subsonic cruise, the rate of fuel consumption should be substantially less, and the combusion exit temperatures should be substantially lower to obtain low specific fuel consumption, possibly in the range of 1700.degree. F. to 1900.degree. F. Under these operating conditions, the engine components which were matched for effective performance at the high power settings with high combustion exit temperatures, are generally not properly matched for efficient operation at the lower power settings of subsonic cruise.
This problem has long been recognized in the prior art, and one possible solution which has been given consideration is the use of a variable area turbine in a supersonic engine. With the variable area turbine, the angle of the stator blades is changed for different operating modes. The blades are moved to a more open position to create a greater turbine nozzle area when the engine is generating greater thrust and thus developing higher combustion exit temperatures, and the stator blades of the turbine area are moved to a more closed position where there is less nozzle area for operation at lower combustion exit temperatures where lower thrust is developed.
This concept of a variable area turbine was analyzed in a paper entitled "Influence of Variable Turbine Geometry of Engine Installation Losses and Cycle Selection", authored by Robert J. May, Jr. and W. F. Zavatkay. This report is designated "AFAPL-TR-73-18" and was presented at a propulsion joint specialists' conference in New Orleans Nov. 27-Dec. 1, 1972. The report concludes that because variable turbine geometry improves the off-design performance, engines incorporating this feature provide good performance over a much broader range of operating conditions. Thus, it is stated in this paper that airplanes designed for a specific mission but incorporating variable turbine geometry engines, will have flexibility to provide good performance for a wide variety of alternate missions.
Three other publications analyzing the variable area turbine concept are the following:
a. "Potential Operating Advantages of a Variable Area Turbine Turbo-Jet", authored by J. W. Ramsay and G. C. Oates, published by the American Society of Mechanical Engineers, United Engineering Center, 345 East 47th St., New York, N.Y. 10017, this publication being contribued for presentation at the Winter Annual Meeting of the Aeorspace Division of the American Society of Mechanical Engineers at New York, Nov. 26-30, 1972. PA1 b. A publication entitled "Experience with a One-Stage Variable Geometry Axial Turbine" by J. Hourmouziadis, K. Hagemeister, O. Rademacher and H. Kolben, Motoren-und Turbinen-Union Munchen GmbH, Dachauer Str. 665, 8000 Munchen 50, Germany. PA1 c. a document entitled "A Paper to be Submitted to the AGARD PROPULSION AND ENERGETICS PANEL, 48th, Meeting (Paris) Sept. 6-10, 1976, VARIABLE GEOMETRY AND MULTICYCLE AERO-ENGINES., authored by R. J. Latimer. PA1 W=Total mass flow rate in lbs. per second PA1 .theta.+=Observed temperature (absolute) divided by standard temperature (518.67.degree. R.) PA1 .delta.+=Observed pressure divided by standard pressure (2116.22 lbs./sq. ft.) PA1 A=Turbine nozzle area. PA1 W=Total mass flow rate in lbs. per second PA1 .theta.+=Observed temperature (absolute) divided by standard temperature (518.67.degree. R.) PA1 .delta.+=Observed pressure divided by standard pressure (2116.22 lbs/sq. ft.)
While these and other analyses have indicated certain operating advantages by use of the variable area turbine, there are still a number of problems in practical implementation of this concept in a jet engine. First, there is a lower peak efficiency due to vane cooling air profile effect and also due to end wall leakage. Also, there is a significant efficiency reduction when the vane area is opened or closed from the design setting. Further, if there is multiple variable-stage turbines, there are rather severe structural mounting problems, and to the best knowledge of the applicant herein, a total satisfactory solution has not been found to this problem. Thus, while the variable area turbine has been demonstated in a test stand, it is still relatively new technology that would likely require large research and development expenditures to bring it to practical production status.
With regard to the prior art disclosed in the patent literature, a number of prior art patents disclose various applications in turbine engines for devices which have passageways directing air from the compressor to bypass the turbine. These various devices are not believed to be directly relevant to the basic concept of the present invention, but they are discussed herein as background information on such turbine bypass apparatus.
U.S. Pat. No. 2,527,732, Imbert, disclosed a turbo-prop engine where in a lower power mode air is directed away from the turbine. When there is requirement for a rapid increase in power, the bypass air is directed into turbine to create the additional power in a relatively short period of time.
U.S. Pat. No. 2,630,673, Woll, disclosed a jet engine where air from the compressor is directed through a bypass passageway to provide cooling for a variable area nozzle at the aft end of the engine.
U.S. Pat. No. 3,049,869, Grenoble, directs air from a low-pressure location in the compressor through a bypass passageway to the aft end of the engine. This bypassed air is combined with over-rich exhaust gas to reburn the mixture at a location rearwardly of the turbine in the engine.
U.S. Pat. No. 3,161,018, Sandre, discloses a combined turbojet-ramjet engine where low pressure air is used in conjunction with a bypass turbojet. U.S. Pat. No. 3,296,800, Keenan et al, shows an arrangement somewhat similar to the Sandre patent noted immediately above.
U.S. Pat. No. 3,514,952, Schumacher et al, discloses a variable bypass turbo-fan engine. During subsonic cruise, the air from the fan is directed through the bypass ducts. During supersonic cruise, valve means close off the bypass ducts so that the air is directed through the compressor and thence to the combustion chamber of the engine.
U.S. Pat. No. 3,520,138, Fox, discloses a plurality of power turbines arranged in series with passageways provided around the second and third turbines, and with valves disposed in the passageways to progressively open or close the passageways. The second and third power turbine combinations are connected to thrust-producing devices for vertical takeoff and landing aircraft or some other desired application.
U.S. Pat. No. 3,641,766, Uehling, discloses an engine arrangement where the thrust output of a gas turbine engine is modulated without the necessity of varying the speed of the engine. This device bypasses a portion of the compressor discharge to a manifold which has a plurality of swirl-inducing nozzles which in turn are able to decrease the thrust output of the engine. The intended result is to decrease the delay time between increased thrust demand and actual thrust output while maintaining engine speed.
U.S. Pat. No. 3,879,941, Sargisson, discloses a variable cycle gas turbine engine with a fan having a forward section axially spaced from an aft section. A variable flow bypassing valve is disposed intermediate the forward and aft fan sections in order that air flow between the forward and aft fan sections may be connected either in series flow relationship or in bypassing parallel relation depending upon the desired mode of engine operation. The variable cycle engine also includes a variable flow geometry inlet duct in direct flow connection to the fan for furnishing an inlet airflow to the fan. Within the variable engine cycle is a core engine having a compressor, combuster and turbine in series flow relationship, wherein the compressor receives a portion of the compressed airflow from the fan. A fan turbine section downwstream of the core engine is also provided to drive the fan.
U.S. Pat. No. 3,841,091, Sargisson, et al discloses a jet engine which is inteded to operate efficiently at both subsonic and supersonic speeds. This embodies a variable cycle tandem propulsion system comprising a forward turbo-fan engine having a fan, gas generator, and power turbine arranged in axially serial flow relation. An independent turbojet engine is co-axially displaced downstream of the turbo-fan engine and includes a compressor, combuster and turbine also arranged in axially spaced serial flow relation. An outer exhaust duct is provided for directing the exhaust stream from the turbo-fan engine rearward around the turbojet engine.
There is also included a variable cross over valve means which may be operated in two modes, e.g., subsonic and supersonic. In the subsonic mode, air flow exiting from the fan which by-passes around the gas generator is directed to the outer exhaust duct means while at the same time a separate inlet ambient air flow stream is directed to the inlet of the turbojet. In its supersonic mode, air flow existing from the fan which by-passes around the gas generator is directed to the turbojet inlet, thereby supercharging the inlet airflow to the turbojet.
U.S. Pat. No. 3,068,644, Worsham et al, relates primarily to a particular type of nozzle wherein shroud flaps are used to control the configuration of a secondary nozzle through which secondary air is directed. U.S. Pat. No. 3,769,797, Stevens, discloses an engine configuration where bypass flow of an engine is used for vertical takeoff and landing mode of operation.
These following patents relate to an air-breathing gas turbine engine where there is valving means for controlling the airflow to "cross over" or be inverted as it passes through the engine: U.S. Pat. Nos. 3,792,584, Klees; 3,854,826, Klees; and 3,938,328, Klees.
The following patents are noted as being generally representative of the state of the art with regard to jet engines in general.
U.S. Pat. No. 2,458,600, Imbert et al, shows an arrangement of a turbo-fan engine.
U.S. Pat. No. 2,505,660, Baumann, discloses a thrust "augmentor" comprising in combination at least two coaxial contrarotationally bladed turbine rotors adapted to be driven by a flow of high velocity combustive gas.
U.S. Pat. No. 3,118,276, Keenen et al, discloses a turbo-fan engine where the fan air communicates with the exhaust gas duct downstream of the turbine or turbines through one or more mixing chutes which extend into the exhaust gas duct.
U.S. Pat. No. 3,316,717, Castle et al, discloses a turbo-fan engine having a variable bypass ratio. This is accomplished by placing fans fore and aft of the gas turbine unit. The fans operate in series for a low bypass ratio, or in parallel for a high bypass ratio.
U.S. Pat. No. 3,903,690, Jones, discloses a turbo-fan engine where all of the bladed stages of the turbine and substantially all of the compressor blades are rotor stages.
U.S. Pat. No. 3,910,375, Hache et al, discloses a jet engine silencer where there is a jet nozzle, and means are provided to inject air into the flow stream emitted from the jet nozzle.
U.S. Pat. No. 3,987,621, Sabatella, Jr. et al, simply discloses a turbo-fan engine where the inner stream includes no noise suppression and the jet exhaust noise generated at take-off is reduced by mechanically suppressing the jet exhaust noise of the outer stream.